System and method for aircraft capacity prediction

ABSTRACT

A technique for identifying, projecting, displaying, and enhancing the thermal capacity for an aircraft is disclosed wherein the thermal capacity is defined as the amount of time or range the aircraft can continue until a thermal limit is exceeded.

CROSS-REFERENCE TO RELATED APPLICATIONS

The present application is a continuation of U.S. Nonprovisional patentapplication Ser. No. 13/589,648 filed Aug. 20, 2012 entitled “System andMethod for Aircraft Thermal Capacity Prediction” and claims the benefitof U.S. Provisional Patent Application Ser. No. 61/525,004 filed Aug.18, 2011 entitled “Aircraft Thermal Capacity Prediction/Adaptation” andwhich are hereby incorporated by reference in their entirety.

GOVERNMENT RIGHTS CLAUSE

This invention was made with Government support under Contract NumberN68335-11-C-0095 awarded by the U.S. Navy. The United States Navy hascertain rights in the invention.

TECHNICAL FIELD OF THE DISCLOSURE

The present disclosure generally relates to systems and methods forprediction of the performance of a physical system and, moreparticularly, to systems and methods for aircraft thermal capacityprediction.

BACKGROUND OF THE DISCLOSURE

For aviation platforms, achieving the highest performance with thelightest weight systems that exceed the required reliability standardsis paramount. Military aircraft often emphasize mission capability asthe key objective, while commercial aircraft often emphasize SpecificFuel Consumption (SFC) and overall life-cycle costs. Historically,mechanical, hydraulic, and pneumatic drive systems have been used onthese platforms for Environmental Control Systems (ECS), and actuationfor engine and flight systems. With recent advances in high-speedbearing and cooling technologies, high-speed electrical machines are nowable to achieve power densities that are competitive with or better thanthe aforementioned conventional drive systems. New systems are alsoincreasingly required to have intelligent actuation control, such thatthey monitor their own health. All of these demands have led to a trendwhere the conventional mechanical, hydraulic, and pneumatic drivesystems are being replaced with electrical systems. The increased use ofsuch systems brings with it an increased need for thermal managementsystems.

In addition to the “electrification” trend of conventional components,demand for on-board power by conventional electronic systems has beenincreasing. Modern aircraft have increasingly powerful ECS, In FlightEntertainment (IFE), and avionics systems; all adding to the demand foronboard power. This trend in civilian aircraft has been preceded by asimilar surge in the power requirements for military aircraft, whereplatforms are moving to the concept of the More Electric Aircraft (MEA).Increasingly powerful avionics, fly-by-wire, electronic warfare, andradar systems have resulted in huge increases in the demand for onboardpower. Considering the smaller dimensions of these military aircraftplatforms, the trend of onboard power demand on a per-passenger orper-volume basis is even greater for military than for civiliantechnology.

Although most electrical systems used onboard aircraft are designed tobe highly efficient, the sheer magnitude of onboard power demand andunique design aspects of onboard aviation systems lead to considerablethermal management challenges. For example, even if a 1 MW system were95% efficient, a total of 50 kW of heat would have to be rejected fromthe aircraft without exceeding rated temperatures. This posessignificant challenges to the thermal management system, as it has toremove all of the generated waste heat with minimal temperature rise ata minimum of weight and volume.

Significant thermal management is required at all levels of powermanagement, including generation, distribution, and conversion powerelectronics (PE), as well as at the application level. At theapplication level, heat is mostly produced by avionics, or, if present,actuators for flight control surfaces. Efficient thermal management isalso required for digital electronics, power electronics andEnvironmental Control Systems (ECS). It is not uncommon for thesesystems to be cascaded; meaning that heat rejected from one system isadded to the heat load of a secondary system before being rejected to anultimate heat sink such as fuel or air.

In addition to the challenge of moving the heat efficiently to theconvective surface, the ultimate rejection of heat is a challenge foraviation platforms. Airborne platforms have only two heat sinksavailable: fuel and ambient air. Fuel is a convenient sink for severalreasons. A large quantity of fuel is available and must be carried onthe aircraft regardless. Heating fuel prior to it entering the enginecombustor is advantageous to the engine efficiency, although this islimited by the thermal stability of conventional jet fuel which, ifcompromised, can foul heat transfer surfaces. Therefore, it is generallypreferred that thermal losses associated with electrification and anyother unwanted heat sources are rejected to the fuel. There areexceptions, for instance, if a component sits in the area of an aircraftwhere the ambient air can effectively accept the losses without the needof a scoop or other component that increases aerodynamic drag. In thatcase, there would not be a need to route fuel to that area.

The More Electric Aircraft (MEA) concept has pushed the use of fuel as aheat sink to the limit. For some short missions on military aircraft,the amount of fuel that will be carried is determined by the electricalheat load rejection capacity requirements rather than the estimatedengine fuel consumption.

While there is a large amount of fuel on-board an aircraft, its use isnot evenly distributed over the flight envelope. During ground idle andidle-descent the fuel flow is very low. During take-off it is extremelyhigh. Thus the fuel flow to the combustor nozzles rarely matches theelectrical loss removal demands required by an MEA. For instance, anaircraft electrical system often requires significant cooling duringidle-descent when electrical loads are relatively high (e.g., fromactuation of flight control surfaces), but fuel flow is extremely low.To meet the cooling requirements, it is often required that the fuel iscirculated back into the fuel tank after being used for cooling. Thisreturn-to-tank arrangement is common on military platforms. Using thefuel tank for thermal energy storage is convenient, but also has itslimitations. At the beginning of the mission the hot fuel returning totank does not significantly increase the overall temperature of the fuelin the tank because of the large thermal mass available. As the missionprogresses and fuel levels are reduced, the high temperature return fuelhas an increasingly greater likelihood of raising the temperature of thefuel in the tank.

The other ultimate heat sink is the ambient air. Air is abundantlyavailable around the aircraft. The quality of this air for use as a heatsink varies widely. On the ground, air can be extremely cold or hot. Theair density at 2,700 meters is about two-thirds that at sea level, andat 5,500 meters the air density is one-half that at sea level. Althoughair at high altitude is cold, viscous heating in the boundary layerbetween the air and the aircraft can be significant, especially atsupersonic velocities. These factors combined make air a much lesseffective cooling fluid than fuel at high altitude, especially at highspeeds.

Exchanging heat with air is also challenging. Air is not a great heattransfer fluid, thus significant heat exchanger surface area is oftenneeded. During flight, air can be scooped from the surrounding aircraftspace and passed through a duct where it interacts with various heatexchangers. Cooling air obtained in this manner is known as “ram air”.Because of the aircraft velocity, little or no fan power is needed todrive the air through the heat exchangers. However, scooping air changesthe aerodynamics of the aircraft and induces drag, resulting in a fuelburn increase. In addition, if the components dumping heat into thisheat sink require cooling during ground idle, a fan or similar fluidpumping device is required to move air across the heat exchangers, orthe heat exchangers must be designed very large to allow for naturalconvection based cooling. In the engine area, heat exchangers can beplaced in the bypass airstream in various ways, including a traditionalheat exchanger directly in the airflow, or a surface cooler inside theengine fan bypass.

As the need to reject heat to the ambient increases with increasinglymore electric aircraft, traditional solutions such as fuel cooling andram air-cooling become problematic. With the transition to compositeskin, fuel-efficient aircraft and increased power demands for aircraftsubsystems such as on-board entertainment systems or advanced radars,significant thermal constraints have arisen wherein the ability of theaircraft to cool critical systems can be compromised during certainflight conditions. For any given aircraft with an established set ofthermal constraints, the inability to sufficiently cool such criticalsystems can result in reduced flight envelopes, component failures, andeven loss of aircraft. Analyzing pre-defined missions provides insightinto potential thermal problems that may or may not develop during amission. Although such insight is beneficial in determining missioncapability for the aircraft, the information quickly becomes obsoleteonce the pilot and/or environmental characteristics deviatesignificantly from the assumed operating conditions originally analyzed.Therefore, missions where a pre-flight analysis indicates no thermalconstraints may in fact result in reduced capability during the missionas a product of unanticipated operating or environmental conditions. Inthe same light, planned mission capability could possibly be extended ifindications of current excess thermal capacity are observed in flight.There therefore exists a need for systems and methods to reliablypredict aircraft thermal capacity during flight.

SUMMARY OF THE DISCLOSURE

Disclosed herein is a technique to identify and project the thermalcapacity for an aircraft either prior to flight or in real-time duringthe flight. A projection of operational capability in response toreal-time and time-history data is invaluable in enhancing thecapability of modern aircraft.

In one embodiment, a method of operating an aircraft is disclosed,comprising the steps of: a) establishing a mission projection for theaircraft; b) flying the aircraft; c) sensing data regarding a currentstate of the aircraft; d) executing a prognostics model using themission projection and the sensed data to estimate a duration until athermal limit of the aircraft is exceeded; e) communicating theestimated duration to an operator of the aircraft; and f) changing anoperating point of the aircraft based upon the estimated duration.

Other embodiments are also disclosed.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic block diagram of one embodiment thermal capacitysystem and method according to the present disclosure;

FIG. 2. is one embodiment of a cockpit display showing real-time thermalcapacity according to one embodiment of the present disclosure;

FIG. 3. is a chart illustrating trade-off between aircraft range andthermal capacity; and

FIG. 4. is an illustration of placement of the display of FIG. 2 in acockpit according to one embodiment of the present disclosure.

DETAILED DESCRIPTION OF THE VARIOUS EMBODIMENTS

For the purpose of promoting an understanding of the principles of theinvention, reference will now be made to the embodiments illustrated inthe drawings and specific language will be used to describe the same. Itwill nevertheless be understood that no limitation of the scope of theinvention is thereby intended. Any alterations and further modificationsin the described embodiments, and any further applications of theprinciples of the invention as described herein are contemplated aswould normally occur to one skilled in the art to which the inventionrelates. One embodiment of the invention is shown in great detail,although it will be apparent to those skilled in the relevant art thatsome features that are not relevant to the present invention may not beshown for the sake of clarity.

Disclosed herein are systems and methods to identify and project thethermal capacity for an aircraft either prior to flight and/or inreal-time during the flight. A projection of operational capability inresponse to real-time and time-history data is invaluable in enhancingthe capability of modern aircraft.

The embodiments disclosed herein have the following distinct advantages:

-   -   1) By quantitatively identifying the current thermal assessment        and thermal range prediction, the pilot/operator can assess the        aircraft's current capabilities and avoid conditions that        compromise the health of the subsystems or of the entire        aircraft.    -   2) By adapting to historic data for a specific aircraft, the        accuracy of the thermal capacity prediction is improved while        not requiring tuning for each specific aircraft.    -   3) By identifying alternative missions/segments/flight        conditions that minimize thermal constraints while maintaining        range and speed, missions/flights that would have been thermally        constrained/aborted are successfully completed.

In the disclosed embodiments, a system and method for real-time thermalcapacity determination is presented that determines the thermallylimited mission time/range through predictive analysis. An exemplarymethod is comprised of:

-   -   1) Mission Projection    -   2) Prognostic Modeling    -   3) Adaptive Techniques    -   4) Thermal Range Prediction    -   5) Identification of Alternative Missions/Segments for Enhanced        Thermal Capacity        A flowchart of one embodiment of the present the thermal        capacity method is shown in FIG. 1. The method uses sensor data        and a predefined mission plan as inputs, and outputs the        thermally limited range/time and alternative missions/segments        for enhanced thermal capacity.        Mission Projection

FIG. 1 schematically illustrates a thermal capacity system and method,indicated generally at 10. In the system 10 of FIG. 1, a predefinedmission 100 is used as an input to develop a mission projection 102. Themission projection 102 is a model of the altitude the aircraft will beflown at for each time segment of the predefined mission 100. Themission projection 102 may be developed in advance of the mission, andthe aircraft altitude, speed, fueling, etc. may be selected in order tokeep the aircraft within its thermal capacity given expectedenvironmental and mission conditions, considering such factors as time,range, mission segments, destination, geographical coordinates, andmeteorological data, to name just a few non-limiting examples. Themission projection 102 is used as an input to the prognostics model 104.

Prognostic Modeling

In the prognostics model 104, a model of the aircraft is utilized toidentify the aircraft's real-time thermal capacitance as a function oftime, based on dynamic internal and external boundary conditions. Itwill be appreciated that the prognostics model 104, as well as the othersystems and methods disclosed herein that require the manipulation ofdata, may be implemented by any type of data processing system capableof manipulating the data in the manner discussed, including one or moreof the following: digital computer, field-programmable gate array(FPGA), complex programmable logic device (CPLD), microprocessor,digital signal processor (DSP), analog computer, and/or any combinationof analog circuitry and/or analog and digital circuitry, to name just afew non-limiting examples. As with any time-domain numerical solution,initial conditions are needed for state variables along with accurateboundary conditions. Sensors 106 on the aircraft measure boundaryconditions (e.g., altitude, Mach number, power level angle (PLA),ambient environment, etc.). Sensors 108 measure state variables (e.g.,fuel temperature, fuel mass, etc.). Using the predefined missionprojection 102, sensor 106 data for boundary conditions, and sensor 108data for initial conditions, the prognostics model 104 predicts whetherand at what point in the mission the temperature limits of the aircraftwill be exceeded.

In one embodiment, the prognostics model 104 includes a boundaryconditions block 110 that accepts the sensor 106 data as inputs. Theoutput of the boundary conditions block 110 is transferred to the heattransfer/fuel flow block 112 and also to the adaptive technique block124 (described below). The output of the heat transfer/fuel flow block112 is summed with the output of the adaptive technique block 118 at114. This summed value is integrated at 116 and this value is providedas input to the fuel temperature/fuel mass block 118. Output from thefuel temperature/fuel mass block 118 is fed back to the heattransfer/fuel flow block 112 and to the adaptive technique block 124.Additionally, the output of the fuel temperature/fuel mass block 118 isused to generate the optimal thermal operation data 120 and the thermalrange prediction 122. The net result of operation of the prognosticsmodel 104 is a calculated range or time within which the aircraft cansafely operate. Predicted aircraft thermal restrictions prior to missioncompletion are determined and relayed to the pilot as a remaining rangeor time until a thermal constraint is exceeded.

Although predefined missions 102 allow for preliminary assessment ofmission capability, uncertainties in flight deviations and environmentalconditions will almost always generate error in this preliminaryanalysis (i.e., cause the actual mission to vary from the projectedmission). To minimize this error, regular refreshes of the prognosticsmodel 104 are performed throughout the flight profile using real-timedata from available sensors 106, 108 to adapt to real-time flightoperation. Meteorological information regarding the ambient environmentis entered into the algorithm prior to takeoff, but if discrepancies arefound between the actual environment and the predicted environment asthe aircraft flies, the algorithm uses actual measured values insubsequent refreshes of the prognostics model 104. In addition, usingfuel temperature/mass sensor 108 data as initial conditions for theprognostics model 104, each refresh of the prognostics model 104 willprovide a more accurate depiction of remaining thermal capacity as themission progresses, including post-refueling. Lastly, the thermalcapacity system 10 also analyzes the current operating point (e.g.,altitude, Mach number, PLA, ambient environment, etc.) to project theremaining thermal capacity of the aircraft, assuming the pilot remainsat the current operating point indefinitely.

Mission Profile Adjustments

The prognostic model 104 can provide an accurate assessment of thermallimitations on the aircraft, but the accuracy of the predictions is afunction of the predicted boundary conditions, namely the assumedmission profile 102. The prognostic model 104 will utilize an assumedMach number and altitude profile to determine boundary conditions suchas solar loading, convection, fuel flow rates, etc., to name just a fewnon-limiting examples. Deviations from the predefined mission 100 willresult in prediction errors with respect to thermal capacity. To addressthis uncertainty, the mission profile adjustment method quantifies thethermal capacity for the current conditions with the assumption that thepilot/operator will return to the predefined mission immediately (i.e.,the pilot is assumed to have only temporarily deviated from thepredefined mission 100).

Adaptive Techniques

Although the aircraft prognostic model 104 captures primary componentperforniance and material properties for an aircraft model, generaltolerances in manufacturing processes and different payloadconfigurations lead to uncertainties in predicted thermal capacity foreach individual aircraft. In addition, as component performance degradesthroughout the useful life of each component on the aircraft, theefficiency (and hence the heat generated by each component) will changeover the life of the aircraft, increasing the discrepancy with theassumed baseline aircraft model. These differences, coupled withmodeling errors, may result in appreciable thermal capacity predictionerror. By employing adaptive or learning techniques, such as neuralnetworks, the thermal capacity system 10 can account for suchuncertainties, providing a more accurate prediction. This is done byproviding the adaptive technique block (with corrective algorithm) 124that receives as inputs the sensor 106 and 108 data, the output of theboundary conditions block 110, and the output of the fueltemperature/fuel mass block 118 (for example).

It will be appreciated that the form of the prognostics model 104 andadaptive technique block (with corrective algorithm) 124 may vary fromapplication to application, based upon the differing characteristics ofthe systems being modeled. For example, the ordinary differentialequation to determine fuel tank temperature is given as,

$\begin{matrix}{{{mC}_{p}\frac{\mathbb{d}T}{\mathbb{d}t}} = {Q_{amb} + Q_{pump} + Q_{in} - Q_{out}}} & (1)\end{matrix}$

The adaptive technique block (with corrective algorithm) 124 adds acorrection term to account for discrepancies between measuredtemperature and predicted temperatures. The corrective term will be afunction of the boundary conditions and temperature error, and will beadded to the equation as,

$\begin{matrix}{{{mC}_{p}\frac{\mathbb{d}T}{\mathbb{d}t}} = {Q_{amb} + Q_{pump} + Q_{in} - Q_{out} - Q_{corr}}} & (2)\end{matrix}$

The correction term is determined by comparing flight data to thepredicted data from the prognostics model 104. The data is run througherror reduction algorithms that analyze multi-dimensional inputs andoutput a corrective term that drives the observed error to zero. Thecorrection term is updated in successive flights as component wearalters the originally predicted performance. It will be appreciated thatthe specific calculation for any corrective term will vary fromapplication to application.

As shown in FIG. 1, the adaptive technique block (with correctivealgorithm) 124 will adjust the heat transfer and/or fuel flow terms fromblock 112 to account for any model discrepancies. In some embodiments,the update rate for the adaptive technique block (with correctivealgorithm) 124 is slower than the regular refresh rates used by theprognostics model 104. The longer time span will allow for a largeraccumulation of data in order to train the adaptive technique block(with corrective algorithm) 124.

Thermal Range Prediction

The output of the thermal capacity system 10 is a thermal rangeprediction 122. The range can be displayed in the form of miles and/ortime remaining before the aircraft exceeds its thermal capacity,providing the pilot/operator with an assessment for probability ofmission completion. FIG. 2 shows one embodiment of a display 200 for thepilot/operator. The display 200 may include a first non-numerical visualindicator, such as the first bar graph 202, that indicates the currentoperating condition of the aircraft with respect to heat rejection. Thebar graph 202 plots in real time the magnitude of the net heat out of(displayed in green, for example) or into (displayed in red, forexample) the aircraft. In the display shown in FIG. 2, the heatrejection off of the aircraft is greater than the heat generatedon-board the aircraft, as indicated by the green bar to the right of themidpoint line 204 (lined for the color green in the patent drawingfigure). The greater the net heat out of the aircraft, the further thegreen bar moves to the right of the midpoint line 204. The midpoint line204 indicates a balance point where the heat rejection off of theaircraft just equals the heat generated onboard the aircraft. Insituations where the heat rejection off of the aircraft is less than theheat generated on-board the aircraft, the bar graph 202 will display ared bar to the left of the midpoint line 204 (not shown). The greaterthe net heat into the aircraft, the further the red bar moves to theleft of the midpoint line 204.

Another feature of the display 200 is the first numerical display 206below the bar graph 202 that indicates the range and/or time (innautical miles and/or minutes, for example) the aircraft can remain inthe current operating point (Mach and altitude, for example) until theaircraft runs out of fuel or exceeds a thermal limit. In someembodiments, the first number display 206 may be color coded. Forexample, if the range and/or time that the current operating point canbe maintained is less than that required for the current missionprojection 102, then the number display 206 may be shown in green. Onthe other hand, showing the number display 206 in red may be used todesignate that the current operating point will be thermally constrainedprior to completion of the mission projection 102.

A further feature that the display 200 may include is a secondnon-numerical visual indicator, such as the second bar graph 208,including a bar indicator 210 that shows the amount of fuel projected tobe remaining on the aircraft at landing. The bar graph 208 may include ared portion 212 (lined for the color red in the patent drawing figure)indicating a range of the amount of fuel on the aircraft that willresult in a thermal constraint being exceeded for the planned mission.Therefore, if the fuel projected to be remaining on the aircraft atlanding (as indicated by the bar indicator 210) is within the redportion 212, a thermal constraint will be exceeded prior to the aircraftlanding and the pilot cannot complete the mission. A second numericaldisplay 214 under the second bar display 208 quantifies the nauticalmiles prior to landing at which the aircraft will exceed the thermalconstraint. In some embodiments, the second number display 206 may becolor coded. For example, if a deficit in the range and/or time requiredfor the current mission projection 102 exists, then the number display214 may be shown in red as a negative number. On the other hand, if nodeficit in the range and/or time required for the current missionprojection 102 exists, the number display 214 may be shown in green as apositive number. The ranges can be displayed in miles and/or time and,in some embodiments, the units may be toggled by the pilot by use of anappropriate control input. The pilot can utilize the information to makedecisions regarding current and future operation of the aircraft toensure successful completion.

Identification of Alternative Missions/Segments/Flight Conditions

The thermal capacity prediction system 10 also include a technique toidentify alternative missions/segments/flight conditions to maximize therange/time before thermal constraints are exceeded, while attempting tomaintain fuel range and aircraft capabilities. One of the outputs of theprognostics model 104 is the optimal thermal operation information 120.The alternative flight plans may be displayed to the pilot/operator inreal-time. In some embodiments, the alternative flight plans aredisplayed as trades between thermal capacity, system capabilities, fuelrange, and aircraft speed. For example, as shown in FIG. 3, the optimalthermal operation information 120 may include information showing therelationship between heat rejected from the aircraft and predicted rangeof the aircraft. A Pareto Front 300 may be indicated on the graph toshow the best compromise between aircraft fuel range and heat rejected.In some embodiments, this optimal point is represented as a Mach numberand altitude at which the pilot should operate the aircraft in order tomaximize heat rejection versus fuel range. Throughout the flight, thepilot will have access to the optimal Mach number and altitude via thepilot vehicle interface and can therefore choose to alter the mission tofly at the preferred operating point

Thermal Capacity Display within the Cockpit

FIG. 4 illustrates one possible location for the thermal capacitydisplay 200 in an aircraft cockpit. While the invention has beenpresented in the context of a specific embodiment, this is for thepurpose of illustration rather than limitation. Other variations andmodifications of the specific embodiment shown and described will beapparent to those skilled in the art within the intended spirit andscope of the invention as set forth in the appended claims.

While the invention has been illustrated and described in detail in thedrawings and foregoing description, the same is to be considered asillustrative and not restrictive in character, it being understood thatonly the preferred embodiments have been shown and described and thatall changes and modifications that come within the spirit of theinvention are desired to be protected. It is also contemplated thatstructures and features embodied in the present examples can be altered,rearranged, substituted, deleted, duplicated, combined, or added to eachother. The articles “the”, “a” and “an” are not necessarily limited tomean only one, but rather are inclusive and open ended so as to include,optionally, multiple such elements.

What is claimed:
 1. A method of operating an aircraft, comprising thesteps of: a) establishing a mission projection for the aircraft; b)flying the aircraft; c) sensing data regarding a current state of theaircraft; d) executing a prognostics model using the mission projectionand the sensed data to estimate a duration until a thermal limit of theaircraft is exceeded, wherein the prognostics model is executedrepeatedly at a first rate, the prognostics model further creating acorrective term at a second rate, wherein the second rate is differentthan the first rate, wherein the duration is estimated based on thecorrective term, wherein the prognostics model outputs optimal thermaloperation information for the aircraft comprising data showing arelationship between heat rejected from the aircraft and predicted rangeof the aircraft; e) communicating the estimated duration to an operatorof the aircraft; and f) changing an operating point of the aircraftbased upon the estimated duration.
 2. The method of claim 1, wherein themission projection comprises a model of an altitude the aircraft will beflown at for each time segment of a predefined mission.
 3. The method ofclaim 1, wherein step (c) comprising sensing data selected from thegroup consisting of: altitude, speed, power level angle, ambienttemperature, ambient pressure, fuel temperature, and fuel mass.
 4. Themethod of claim 1, wherein the prognostics model models the aircraft'sreal-time thermal capacitance as a function of time based on dynamicinternal and external boundary conditions.
 5. The method of claim 1,wherein the prognostics model comprises: a boundary conditions blockthat accepts the sensed data and mission projection as inputs andcreates a boundary conditions output; an adaptive technique section thataccepts the sensed data, the boundary conditions output, and a fueltemperature/fuel mass output as inputs and creates the corrective termas an output; a heat transfer/fuel flow block that accepts the boundaryconditions output and the fuel temperature/fuel mass output as inputsand creates a heat transfer/fuel flow output; a summing block that sumsthe heat transfer/fuel flow output and the corrective term to create asum output; an integrator that integrates the sum output to create anintegrator output; and a fuel temperature/fuel mass block that acceptsas an input the integrator output and produces the fuel temperature/fuelmass output.
 6. The method of claim 5, wherein the fuel temperature/fuelmass output is used to generate the estimated duration.
 7. The method ofclaim 5, wherein the boundary conditions output is selected from thegroup consisting of: solar loading, convection, and fuel flow rate. 8.The method of claim 5, wherein the prognostics model is executedrepeatedly at the first rate while the aircraft is flown.
 9. The methodof claim 8, wherein the adaptive technique creates the corrective termat the second rate, wherein the second rate is slower than the firstrate.
 10. The method of claim 1, wherein the operating point is selectedfrom the group consisting of: altitude, speed, and fueling.
 11. Themethod of claim 1, wherein if the aircraft is not currently on themission projection, the prognostics model estimates the duration byassuming that the operator will return the aircraft to the missionprojection immediately.
 12. The method of claim 1, wherein the estimatedduration comprises data selected from the group consisting of: distance,and time.
 13. The method of claim 1, wherein step (e) comprises creatinga display visible to the operator, the display comprising: a firstnon-numerical visual indicator communicating a current of net heat intoor out of the aircraft; a first numerical display communicating theestimated duration; a second non-numerical visual indicatorcommunicating an amount of fuel projected to be remaining when theaircraft lands; and a second numerical display communicating a distanceprior to landing that the aircraft will exceed the thermal limit. 14.The method of claim 1, wherein the optimal thermal operation informationincludes a Pareto front showing a best compromise between heat rejectedfrom the aircraft and predicted range of the aircraft.
 15. The method ofclaim 1, wherein steps (a), (d) and (e) are performed by a processingdevice selected from the group consisting of: a digital computer, afield-programmable gate array, a complex programmable logic device, amicroprocessor, a digital signal processor, an analog computer, analogcircuitry, and digital circuitry.
 16. The method of claim 1, wherein themission projection is based on one or more factors selected from thegroup consisting of: time, range, mission segments, destination,geographical coordinates, or meteorological data.
 17. The method ofclaim 1, wherein the optimal thermal operation information comprises analternative flight plan.
 18. The method of claim 17, wherein thealternative flight plan is displayed in real-time.
 19. The method ofclaim 17, wherein the alternative flight plan is displayed as tradesbetween thermal capacity, system capabilities, fuel range, and aircraftspeed.
 20. The method of claim 17, wherein the alternative flight planis displayed as Mach number and altitude.